Composite airfoil assembly with separate airfoil, inner band, and outer band

ABSTRACT

Airfoil assemblies for gas turbine engines are provided. For example, an airfoil assembly comprises an airfoil, an inner band defining an inner opening shaped complementary to an inner end of the airfoil, and an outer band defining an outer opening shaped complementary to an outer end of the airfoil. The airfoil inner end is received with the inner opening, and the airfoil outer end is received within the outer opening. A strut extends radially through an airfoil cavity. A first pad is defined at a first radial location within the cavity. A second pad is defined within the cavity at a second, different radial location. In some embodiments, the airfoil assembly inner band includes a first inner flange, through which the inner band is secured to a support structure, and the outer band includes a first outer flange, through which the outer band is secured to a support structure.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of and claims priority to U.S.application Ser. No. 15/922,263, filed Mar. 15, 2018, the contents ofwhich are incorporated herein by reference.

FEDERALLY SPONSORED RESEARCH

This invention was made with government support under contact numberFA8650-15-D-2501 awarded by the Department of the Air Force. The U.S.government may have certain rights in the invention.

FIELD

The present subject matter relates generally to gas turbine engines.More particularly, the present subject matter relates to compositeairfoil assemblies for gas turbine engines, such as composite turbinenozzle fairings for gas turbine engines.

BACKGROUND

More commonly, non-traditional high temperature composite materials,such as ceramic matrix composite (CMC) materials, are being used inapplications such as gas turbine engines. Components fabricated fromsuch materials have a higher temperature capability compared withtypical components, e.g., metal components, which may allow improvedcomponent performance and/or increased engine temperatures. Compositecomponents may provide other advantages as well, such as an improvedstrength to weight ratio.

Typically, a CMC turbine nozzle fairing comprises an airfoil, an innerband, and an outer band that are integrally formed as one singlecomponent, with curved transition zones between the airfoil and each ofthe inner band and outer band. However, the transition from the airfoilto the band sections in the CMC turbine nozzle fairing generallycomprises complex shapes in the vicinity of the curvature such that thenozzle fairings are difficult to lay up, resulting in a longmanufacturing cycle time and low yield, and also are difficult tocompact, often resulting in poor compaction. Additionally, thermaldifferences, i.e., a thermal fight, between the airfoil and bandsproduce high stresses in the nozzle fairings, which limits theacceptability of part defects and results in tighter inspection limitsfor non-destructive examination of the parts. Moreover, known CMC nozzlefairings typically are singlets and can allow leakage between eachseparate nozzle fairing.

Accordingly, improved airfoil assemblies would be useful. In particular,an airfoil assembly comprising an airfoil that is separate from each ofthe inner band and outer band would be advantageous. Further, an airfoilassembly having a separate airfoil, inner band, and outer band that issimply supported, with a positively located airfoil, would be desirable.

BRIEF DESCRIPTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

In one exemplary embodiment of the present subject matter, an airfoilassembly for a gas turbine engine is provided. The airfoil assemblycomprises an airfoil having a concave pressure side opposite a convexsuction side and an inner end radially spaced apart from an outer end.The pressure side and the suction side extend axially from a leadingedge to a trailing edge. The airfoil assembly further comprises an innerband defining an inner opening shaped complementary to the inner end ofthe airfoil and an outer band defining an outer opening shapedcomplementary to the outer end of the airfoil. The inner end of theairfoil is received with the inner opening and the outer end of theairfoil is received within the outer opening. The airfoil assembly alsocomprises a strut extending radially through a cavity defined by theairfoil, as well as a first pad defined at a first radial locationwithin the cavity and a second pad defined at a second radial locationwithin the cavity. The first radial location is different from thesecond radial location.

In another exemplary embodiment of the present subject matter, anairfoil assembly for a gas turbine engine is provided. The airfoilassembly comprises an airfoil having a concave pressure side opposite aconvex suction side and an inner end radially spaced apart from an outerend. The pressure side and the suction side extend axially from aleading edge to a trailing edge. The airfoil assembly also comprises aninner band defining an inner opening shaped complementary to the innerend of the airfoil and an outer band defining an outer opening shapedcomplementary to the outer end of the airfoil. The inner end of theairfoil is received with the inner opening and the outer end of theairfoil is received within the outer opening. The inner band includes afirst flowpath surface, a first non-flowpath surface opposite the firstflowpath surface, and a first inner flange extending radially from thefirst non-flowpath surface. The outer band includes a second flowpathsurface, a second non-flowpath surface opposite the second flowpathsurface, and a first outer flange extending radially from the secondnon-flowpath surface. The inner band is secured to an inner supportstructure by a first inner fastener extending through the first innerflange, and the outer band is secured to an outer support structure by afirst outer fastener extending through the first outer flange.

In a further exemplary embodiment of the present subject matter, anairfoil assembly for a gas turbine engine is provided. The airfoilassembly comprises an airfoil having a concave pressure side opposite aconvex suction side and an inner end radially spaced apart from an outerend. The pressure side and the suction side extend axially from aleading edge to a trailing edge. The airfoil assembly further comprisesan inner band defining an inner opening shaped complementary to theinner end of the airfoil and an outer band defining an outer openingshaped complementary to the outer end of the airfoil. The inner end ofthe airfoil is received with the inner opening and the outer end of theairfoil is received within the outer opening. The inner band includes afirst flowpath surface, a first non-flowpath surface opposite the firstflowpath surface, and a first inner flange extending radially from thefirst non-flowpath surface. The outer band includes a second flowpathsurface, a second non-flowpath surface opposite the second flowpathsurface, and a first outer flange extending radially from the secondnon-flowpath surface. The inner band is secured to an inner supportstructure by a first inner fastener extending through the first innerflange, and the outer band is secured to an outer support structure by afirst outer fastener extending through the first outer flange. Moreover,a strut extends radially through a cavity defined by the airfoil, and afirst pad is defined at a first radial location within the cavity and asecond pad is defined at a second radial location within the cavity. Thefirst radial location is different from the second radial location. Eachof the inner band, outer band, and airfoil are formed from a ceramicmatrix composite material.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 provides a schematic cross-section view of an exemplary gasturbine engine according to various embodiments of the present subjectmatter.

FIG. 2 provides a perspective view of a doublet airfoil assembly, havingtwo airfoils separate from an inner band and outer band, according to anexemplary embodiment of the present subject matter.

FIG. 3 provides an axial cross-section view of one airfoil of theexemplary airfoil assembly of FIG. 2.

FIG. 4 provides a side schematic view of an airfoil assembly having asingle pinned flange on each of the inner band and outer band, accordingto an exemplary embodiment of the present subject matter.

FIG. 5 provides a side schematic view of an airfoil assembly having adouble pinned flange on each of the inner band and outer band, accordingto an exemplary embodiment of the present subject matter.

DETAILED DESCRIPTION

Reference will now be made in detail to present embodiments of theinvention, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the invention.

As used herein, the terms “first”, “second”, and “third” may be usedinterchangeably to distinguish one component from another and are notintended to signify location or importance of the individual components.

The terms “forward” and “aft” refer to relative positions within a gasturbine engine or vehicle, and refer to the normal operational attitudeof the gas turbine engine or vehicle. For example, with regard to a gasturbine engine, forward refers to a position closer to an engine inletand aft refers to a position closer to an engine nozzle or exhaust.

The terms “upstream” and “downstream” refer to the relative directionwith respect to fluid flow in a fluid pathway. For example, “upstream”refers to the direction from which the fluid flows, and “downstream”refers to the direction to which the fluid flows.

The terms “coupled,” “fixed,” “attached to,” and the like refer to bothdirect coupling, fixing, or attaching, as well as indirect coupling,fixing, or attaching through one or more intermediate components orfeatures, unless otherwise specified herein.

The singular forms “a”, “an”, and “the” include plural references unlessthe context clearly dictates otherwise.

Approximating language, as used herein throughout the specification andclaims, is applied to modify any quantitative representation that couldpermissibly vary without resulting in a change in the basic function towhich it is related. Accordingly, a value modified by a term or terms,such as “about”, “approximately”, and “substantially”, are not to belimited to the precise value specified. In at least some instances, theapproximating language may correspond to the precision of an instrumentfor measuring the value, or the precision of the methods or machines forconstructing or manufacturing the components and/or systems. Forexample, the approximating language may refer to being within a 10percent margin.

Here and throughout the specification and claims, range limitations arecombined and interchanged, such ranges are identified and include allthe sub-ranges contained therein unless context or language indicatesotherwise. For example, all ranges disclosed herein are inclusive of theendpoints, and the endpoints are independently combinable with eachother.

Referring now to the drawings, wherein identical numerals indicate thesame elements throughout the figures, FIG. 1 is a schematiccross-sectional view of a gas turbine engine in accordance with anexemplary embodiment of the present disclosure. More particularly, forthe embodiment of FIG. 1, the gas turbine engine is a high-bypassturbofan jet engine 10, referred to herein as “turbofan engine 10.” Asshown in FIG. 1, the turbofan engine 10 defines an axial direction A(extending parallel to an axial centerline 12 provided for reference)and a radial direction R. In general, the turbofan 10 includes a fansection 14 and a core turbine engine 16 disposed downstream from the fansection 14.

The exemplary core turbine engine 16 depicted generally includes asubstantially tubular outer casing 18 that defines an annular inlet 20.The outer casing 18 encases, in serial flow relationship, a compressorsection including a booster or low pressure (LP) compressor 22 and ahigh pressure (HP) compressor 24; a combustion section 26; a turbinesection including a high pressure (HP) turbine 28 and a low pressure(LP) turbine 30; and a jet exhaust nozzle section 32. A high pressure(HP) shaft or spool 34 drivingly connects the HP turbine 28 to the HPcompressor 24. A low pressure (LP) shaft or spool 36 drivingly connectsthe LP turbine 30 to the LP compressor 22.

For the depicted embodiment, fan section 14 includes a fan 38 having aplurality of fan blades 40 coupled to a disk 42 in a spaced apartmanner. As depicted, fan blades 40 extend outward from disk 42 generallyalong the radial direction R. The fan blades 40 and disk 42 are togetherrotatable about the axial centerline 12 by LP shaft 36. In someembodiments, a power gear box having a plurality of gears may beincluded for stepping down the rotational speed of the LP shaft 36 to amore efficient rotational fan speed.

Referring still to the exemplary embodiment of FIG. 1, disk 42 iscovered by rotatable front nacelle 48 aerodynamically contoured topromote an airflow through the plurality of fan blades 40. Additionally,the exemplary fan section 14 includes an annular fan casing or outernacelle 50 that circumferentially surrounds the fan 38 and/or at least aportion of the core turbine engine 16. It should be appreciated thatnacelle 50 may be configured to be supported relative to the coreturbine engine 16 by a plurality of circumferentially-spaced outletguide vanes 52. Moreover, a downstream section 54 of the nacelle 50 mayextend over an outer portion of the core turbine engine 16 so as todefine a bypass airflow passage 56 therebetween.

During operation of the turbofan engine 10, a volume of air 58 entersturbofan 10 through an associated inlet 60 of the nacelle 50 and/or fansection 14. As the volume of air 58 passes across fan blades 40, a firstportion of the air 58 as indicated by arrows 62 is directed or routedinto the bypass airflow passage 56 and a second portion of the air 58 asindicated by arrows 64 is directed or routed into the LP compressor 22.The ratio between the first portion of air 62 and the second portion ofair 64 is commonly known as a bypass ratio. The pressure of the secondportion of air 64 is then increased as it is routed through the highpressure (HP) compressor 24 and into the combustion section 26, where itis mixed with fuel and burned to provide combustion gases 66.

The combustion gases 66 are routed through the HP turbine 28 where aportion of thermal and/or kinetic energy from the combustion gases 66 isextracted via sequential stages of HP turbine stator vanes 68 that arecoupled to the outer casing 18 and HP turbine rotor blades 70 that arecoupled to the HP shaft or spool 34, thus causing the HP shaft or spool34 to rotate, thereby supporting operation of the HP compressor 24. Thecombustion gases 66 are then routed through the LP turbine 30 where asecond portion of thermal and kinetic energy is extracted from thecombustion gases 66 via sequential stages of LP turbine stator vanes 72that are coupled to the outer casing 18 and LP turbine rotor blades 74that are coupled to the LP shaft or spool 36, thus causing the LP shaftor spool 36 to rotate, thereby supporting operation of the LP compressor22 and/or rotation of the fan 38.

The combustion gases 66 are subsequently routed through the jet exhaustnozzle section 32 of the core turbine engine 16 to provide propulsivethrust. Simultaneously, the pressure of the first portion of air 62 issubstantially increased as the first portion of air 62 is routed throughthe bypass airflow passage 56 before it is exhausted from a fan nozzleexhaust section 76 of the turbofan 10, also providing propulsive thrust.The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section32 at least partially define a hot gas path 78 for routing thecombustion gases 66 through the core turbine engine 16.

In some embodiments, components of turbofan engine 10, particularlycomponents within or defining the hot gas path 78, may comprise acomposite material, such as a ceramic matrix composite (CMC) materialhaving high temperature capability. Composite materials generallycomprise a fibrous reinforcement material embedded in matrix material,e.g., a ceramic matrix material. The reinforcement material serves as aload-bearing constituent of the composite material, while the matrix ofa composite material serves to bind the fibers together and act as themedium by which an externally applied stress is transmitted anddistributed to the fibers.

Exemplary CMC materials may include silicon carbide (SiC), silicon,silica, or alumina matrix materials and combinations thereof. Ceramicfibers may be embedded within the matrix, such as oxidation stablereinforcing fibers including monofilaments like sapphire and siliconcarbide (e.g., Textron's SCS-6), as well as rovings and yarn includingsilicon carbide (e.g., Nippon Carbon's NICALON®, Ube Industries'TYRANNO®, and Dow Corning's SYLRAMIC®), alumina silicates (e.g.,Nextel's 440 and 480), and chopped whiskers and fibers (e.g., Nextel's440 and SAFFIL®), and optionally ceramic particles (e.g., oxides of Si,Al, Zr, Y, and combinations thereof) and inorganic fillers (e.g.,pyrophyllite, wollastonite, mica, talc, kyanite, and montmorillonite).For example, in certain embodiments, bundles of the fibers, which mayinclude a ceramic refractory material coating, are formed as areinforced tape, such as a unidirectional reinforced tape. A pluralityof the tapes may be laid up together (e.g., as plies) to form a preformcomponent. The bundles of fibers may be impregnated with a slurrycomposition prior to forming the preform or after formation of thepreform. The preform may then undergo thermal processing, such as a cureor burn-out to yield a high char residue in the preform, and subsequentchemical processing, such as melt-infiltration with silicon, to arriveat a component formed of a CMC material having a desired chemicalcomposition. In other embodiments, the CMC material may be formed as,e.g., a carbon fiber cloth rather than as a tape.

Turning to FIG. 2, an exemplary airfoil assembly 100, e.g., a nozzlefairing assembly for turbofan engine 10, is illustrated. The airfoilassembly 100 comprises two airfoils 102, an inner band 104, and an outerband 106. Because the airfoil assembly 100 includes two airfoils 102received within a single inner band 104 and a single outer band 106, theassembly may be referred to as a doublet airfoil assembly 100. In thedepicted exemplary embodiment, each airfoil 102, the inner band 104, andthe outer band 106 is formed from a CMC material. As shown in FIG. 2,the exemplary CMC airfoils 102 each include a concave pressure side 108opposite a convex suction side 110. Opposite pressure and suction sides108, 110 of the airfoil 102 radially extend between an inner end 112 andan outer end 114 along an airfoil span S. Moreover, pressure and suctionsides 108, 110 of the airfoil 102 extend axially between a leading edge116 and an opposite trailing edge 118, and the pressure and suctionsides 108, 110 define an outer surface 120 of the airfoil 102. Further,each illustrated airfoil 102 includes an inner parapet 122 that extendsabout the airfoil 102 at its inner end 112, and an outer parapet 124that extends about the airfoil 102 at its outer end 114. Additionally,referring to FIG. 3, each airfoil 102 includes a trailing edge portion126 that defines its trailing edge 118. The trailing edge portion 126 islocated aft of a cavity 128 defined by the airfoil 102. The cavity 128extends the radial length, i.e., the span S, of the airfoil 102.

As further shown in FIG. 2, the inner and outer bands 104, 106 arerelatively thin CMC structures that are separate from the airfoil 102.That is, each of the airfoil 102, inner band 104, and outer band 106 areseparately formed from a CMC material such that each component is anindividual piece. In the depicted embodiment, the airfoil assembly 100is a turbine nozzle fairing assembly, and a plurality of airfoilassemblies 100 may be positioned circumferentially adjacent one anotherto form an annular turbine nozzle stage, e.g., a plurality of turbinenozzles positioned circumferentially about the axial centerline 12 ofthe engine 10. As such, each of the inner band 104 and outer band 106form a liner along the hot gas path 78, protecting metallic componentsand the like from the heat of the combustion gases 66.

As illustrated in FIG. 2, the inner band 104 defines two inner openings130 that are shaped complementary to the inner end 112 of each airfoil102. As such, the inner end 112 of each airfoil 102 is received within arespective inner opening 130. Similarly, the outer band 106 defines twoouter openings 132 shaped complementary to the outer end 114 of eachairfoil 102, such that the outer end 114 of each airfoil 102 is receivedwithin a respective outer opening 132. An inner seal 134 extends aroundthe inner end 112 of each airfoil 102 such that the inner seal 134 ispositioned between the inner end 112 and the inner band 104 to sealagainst leakage through the inner opening 130. Likewise, an outer seal136 extends around the outer end 114 of each airfoil 102 such that theouter seal 136 is positioned between the outer end 114 and the outerband 106 to seal against leakage through the outer opening 132.Moreover, the inner and outer seals 134, 136 are positioned to engagethe inner band 104 and outer band 106, respectively. The inner and outerseals 134, 136 are illustrated for only one airfoil 102 in FIG. 2; theseals 134, 136 for the other airfoil 102 in FIG. 2 are omitted forclarity.

Turning to FIGS. 3, 4, and 5, the airfoil assembly 100 will be describedin greater detail. FIG. 3 provides an axial cross-section view of oneairfoil 102 of the exemplary airfoil assembly 100 of FIG. 2. FIG. 4provides a side schematic view of the airfoil assembly 100 having asingle pinned flange on each of the inner band 104 and outer band 106,according to one exemplary embodiment of the present subject matter.FIG. 5 provides a side schematic view of the airfoil assembly 100 havinga double pinned flange on each of the inner band 104 and outer band 106,according to another exemplary embodiment of the present subject matter.

As illustrated in FIGS. 3 and 5, a strut 140 extends radially througheach airfoil 102, the inner band 104, and the outer band 106 of theairfoil assembly 100. More particularly, the strut 140 extends througheach airfoil 102 within the cavity 128 defined in the airfoil 102. Thestrut 140 includes a first pad 142 at a first radial location R1 withinthe cavity 128 and a second pad 144 at a second radial location R2within the cavity 128. As depicted in FIG. 5, the first radial locationR1 is different from the second radial location R2, and the first andsecond radial locations R1, R2 are determined from the axial centerline12 of the engine 10. In other embodiments, the first and second pads142, 144 may be defined on the airfoil 102 rather than the strut 140.Although not illustrated in FIG. 4, it will be appreciated that a strut140 could extend through the airfoil 102 as shown in FIGS. 3 and 5.

Keeping with FIGS. 3 and 5, the airfoil 102 is also constrained axiallysuch that axial loading of the airfoil 102 transfers the load to aninner support structure 174 and an outer support structure 178. Asshown, a radially extending first slot 146 is defined in the trailingedge portion 126 of each airfoil 102. Each first slot 146 is configuredfor receipt of a first pin 148. Further, a radially extending secondslot 150 is defined in each airfoil 102 for receipt of a second pin 152.The depicted airfoil assembly 100 could be configured as a first stageor a second stage turbine nozzle assembly, i.e., when installed within agas turbine engine. If configured as a second stage assembly 100, thefirst slot 146 is defined in the outer end 114 of each airfoil 102, andthe second slot 150 is defined in the inner end 112 of each airfoil 102.A first aperture 154 is defined in the outer support structure 178adjacent each first slot 146 such that each first pin 148 passes throughthe first aperture 154 into the first slot 146, and a second aperture156 is defined in the inner support structure 174 adjacent each secondslot 150 such that each second pin 152 passes through the secondaperture 156 into the second slot 150. Thus, each airfoil 102 isconstrained axially by a pair of pins 148, 152 extending radially intothe airfoil 102, the first pin 148 at the outer end 114 of the trailingedge portion 126, and the second pin 152 at the inner end 112 of theairfoil 102 just aft of the cavity 128. Further, the pins 148, 152prevent the airfoil 102 from bottoming out within the openings 130, 132,as well as from pulling out of the openings 130, 132.

For a second stage nozzle assembly, the depicted pin configurationsupports and locates the respective airfoil 102, with the axial loadspassing into the inner and outer support structures 174, 178. However,it will be appreciated that where the airfoil assembly 100 is a firststage nozzle assembly (e.g., as illustrated in FIG. 4), the first slots146 are typically defined in the inner end 112 of the airfoil 102, andthe second slots 150 are typically defined in the outer end 114 of theairfoil 102. Likewise, in such embodiments, the first apertures 154 aredefined in the inner support structure 174 such that the first pins 148pass through the first aperture 154 into the first slot 146 at the innerend 112 of the trailing edge portion 126 of each airfoil 102, and thesecond apertures 156 are defined in the outer support structure 178 suchthat the second pins 152 pass through the outer support structure 178into the second slot 150 just aft of the cavity 128 of each airfoil 102.Such a pin configuration adequately axially constrains the first stagenozzle assembly 100, supporting and locating each airfoil 102 of theassembly 100 such that the axial loads pass into the inner and outersupport structures 174, 178. As such, the second stage nozzle iscantilevered from the outer casing 18, with an inter-stage seal attachedto the inner support structure 174. The cantilevered configuration maybe applied to nozzles of any turbine stage that requires the nozzles tobe cantilevered, e.g., third, fourth, etc. stage nozzles. In contrast,the first stage nozzle is non-cantilevered, with the inner and outersupport structures 174, 178 being supported by the engine's staticstructure, e.g., outer casing 18.

Further, in either the first or second stage embodiment of the airfoilassembly 100, the first and second pins 148, 152 radially constrain eachairfoil 102. Moreover, referring to FIG. 3, each airfoil 102 isconfigured to contact the first and second pads 142, 144 of the strut140 extending through the airfoil 102 and load the slot 146 into the pin148 when the airfoil 102 is tangentially loaded. More specifically, wheneach airfoil 102 experiences a tangential load, the airfoil 102 isconfigured to load into the pads 142, 144 such that an interior surface138 of the airfoil 102, which defines the cavity 128 of the airfoil 102,contacts the first and second pads 142, 144. More generally, e.g., forembodiments in which the pads 142, 144 are not part of or attached tothe strut 140, the first and second pads 142, 144 provide a loadingsurface or stop for the airfoil 102 to load into the strut 140 when theairfoil 102 is tangentially loaded. As such, the first and second pads142, 144 and the first pin 148 transfer the tangential loads to thestrut 140 and inner support structure 174. The first and second pins148, 152 support and locate each airfoil 102 while transferring radialand axial loads to the inner and outer support structures 174, 178.

Further, in exemplary embodiments of the airfoil assembly 100, the strut140 is formed from a metallic material, such as a metal or metal alloy.Accordingly, the CMC airfoil 102 and the metallic strut 140 havedifferent coefficients of thermal expansion a. As previously described,each airfoil 102 is separate from the inner band 104 and outer band 106,and is not attached or fastened to the strut 140, and as such, theairfoil 102 is free to float radially to accommodate the difference incoefficients of thermal expansion a between the airfoil 102 and thestrut 140. That is, the strut 140 will begin to expand at a lowertemperature than each airfoil 102, and the airfoil 102 has freedom tomove radially to accommodate the thermal expansion of the strut 140.

Turning now to FIGS. 4 and 5, the inner band 104 includes a firstflowpath surface 160 and a first non-flowpath surface 162 opposite thefirst flowpath surface 160. Similarly, the outer band 106 includes asecond flowpath surface 164 and a second non-flowpath surface 166opposite the second flowpath surface 164. The first and second flowpathsurfaces 160, 164 help define the hot gas path 78, while the first andsecond non-flowpath surfaces 162, 166 are positioned outside of the hotgas path 78.

Referring particularly to FIG. 4, in one exemplary embodiment, a firstinner flange 168 extends radially from the first non-flowpath surface162, and a first outer flange 170 extends radially from the secondnon-flowpath surface 166. A first inner fastener 172 extends through anaperture 173 in the first inner flange 168 and into the inner supportstructure 174 to secure the inner band 104 to the inner supportstructure 174. Likewise, a first outer fastener 176 extends through anaperture 177 in the first outer flange 170 and into the outer supportstructure 178 to secure the outer band 106 to the outer supportstructure 178. Thus, the exemplary embodiment of FIG. 4 comprises asingle pinned flange on each of the inner band 104 and outer band 106 toattach the bands 104, 106 to their respective support structures 174,178, e.g., metallic hangers or the like. It will be appreciated thatmore than one inner fastener 172 and outer fastener 176, e.g., eachinner fastener 172 and each outer fastener 176 circumferentially spacedapart from one another, may pass through the inner band 104 and outerband 106, respectively, to secure the inner and outer bands 104, 106 totheir respective support structures 174, 178. In such embodiments, theaperture 173 for an additional inner fastener 172 and the aperture 177for an additional outer fastener 176 extend in the circumferentialdirection C to account for the thermal expansion difference between thesupport structures 174, 178 (e.g., metal hangers) and the CMC bands 104,106.

As shown in FIG. 4, the first inner flange 168 extends from the firstnon-flowpath surface 162 near a mid-portion 180 of the inner band 104,and the first outer flange 170 extends from the second non-flowpathsurface 166 near an aft end 182 of the outer band 106. Preferably, eachof the first inner flange 168 and the first outer flange 170 are definedon the respective band 104, 106 near a resultant pressure force tominimize moment into the pin joint between the first inner flange 168and the inner support structure 174 and the first outer flange 170 andthe outer support structure 178. That is, the various forces formed by apressure distribution across the bands 104, 106 produce a moment at eachpin joint. The position of each flange 168, 178 on its respective band104, 106 is selected to minimize the moment created by the pressureforces that is taken through each pin 148, 152.

Referring now to FIG. 5, in another exemplary embodiment of the airfoilassembly 100, each of the inner band 104 and outer band 106 includes adouble pinned flange attaching the band 104, 106 to its respectivesupport structure 174, 178. More particularly, in addition to the firstinner flange 168, the inner band 104 includes a second inner flange 184extending radially from the first non-flowpath surface 162, and inaddition to the first outer flange 170, the outer band 106 includes asecond outer flange 186 extending radially from the second non-flowpathsurface 166. A second inner fastener 188 extends through an aperture 189(FIG. 2) in the second inner flange 184 and into the inner supportstructure 174, and a second outer fastener 190 extends through anaperture 191 (FIG. 2) in the second outer flange 186 and into the outersupport structure 178. As such, the first and second inner fasteners172, 188 secure the inner band 104 to the inner support structure 174,and the first and second outer fasteners 176, 190 secure the outer band106 to the outer support structure 178. As described above, more thanone inner fastener 188 and outer fastener 190, e.g., each inner fastener188 and each outer fastener 190 circumferentially spaced apart from oneanother, may pass through the inner band 104 and outer band 106,respectively, to secure the inner and outer bands 104, 106 to theirrespective support structures 174, 178. In such embodiments, theaperture 189 for an additional inner fastener 188 and the aperture 191for an additional outer fastener 190 extend in the circumferentialdirection C to account for the thermal expansion difference between thesupport structures 174, 178 (e.g., metal hangers) and the CMC bands 104,106.

As further depicted in FIG. 5, in the double pinned embodiment, thefirst inner flange 168 extends from the first non-flowpath surface 162of the inner band 104 near the mid-portion 180 of the inner band 104.The second inner flange 184 extends from the first non-flowpath surface162 near a forward end 192 of the inner band 104. Moreover, the firstouter flange 170 extends from the second non-flowpath surface 166 of theouter band 106 near the aft end 182 of the outer band 106, while thesecond outer flange 186 extends from the second non-flowpath surface 166near a forward end 194 of the outer band 106. As a result, the firstinner flange 168 and the second inner flange 184 extend from the firstnon-flowpath surface 162 of the inner band 104 such that the first andsecond inner flanges 168, 184 straddle the inner support structure 174.Similarly, the first outer flange 170 and the second outer flange 186extend from the second non-flowpath surface 166 of the outer band 106such that the first and second outer flanges straddle the outer supportstructure 178. As such, the first and second inner flanges 168, 184 wraparound the inner support structure 174 and the first and second outerflanges 170, 186 wrap around the outer support structure 178, therebyforming a CMC liner around each of the first and outer supportstructures 174, 178 out of the direct flow path 78 of the combustiongases 66. As previously described, the support structures 174, 178 maybe hangers formed from a metallic material, e.g., a metal or metalalloy, such that the CMC bands 104, 106 with their flanges 168, 184,170, 186 help protect the support structures 174, 178 from the hightemperature combustion gases 66.

Additionally, although the airfoil assembly 100 depicted in theexemplary embodiments includes two airfoils, the airfoil assembly 100described herein could be formed as a singlet, triplet, etc. Moreparticularly, for a singlet assembly 100, the inner band 104 defines oneinner opening 130 and the outer band 106 defines one outer opening 132.A single airfoil 102 extends from the inner band 104 to the outer band106 with the inner end 112 of the airfoil 102 positioned in the inneropening 130 and the outer end 114 of the airfoil 102 positioned in thecorresponding outer opening 132. A triplet airfoil assembly 100 wouldinclude three airfoils 102, with each airfoil 102 extending between aninner opening 130 and an outer opening 132 such that the tripletassembly 100 includes three inner openings 130 and three outer openings132. An airfoil assembly 100 including any appropriate number ofairfoils 102 extending from the inner band 104 and outer band 106 may beformed. It will be appreciated that, by reducing the number of inner andouter band segments, split line leakage, i.e., leakage between adjacentinner and outer band segments, can be reduced. More specifically, splitline leakage is eliminated where splits in the inner band 104 and outerband 106 are eliminated to form a doublet, triplet, etc. airfoilassembly 100. However, as shown in FIGS. 2, 4, and 5, seals 196 also maybe included between adjacent inner and outer bands 104, 106 to helpminimize leakage between band segments.

As described herein, the airfoils 102, inner band 104, and outer band106 may be formed from a CMC material. After laying up the CMC plies aspreviously described to form a layup or preforms for each of theairfoils 102, inner band 104, and outer band 106, the layups aredebulked and, if appropriate, cured while subjected to elevatedpressures and temperatures to produce cured preforms, e.g., the layupsmay be cured in an autoclave to form autoclaved bodies. In exemplaryembodiments, the autoclaved bodies are then heated (fired) in a vacuumor inert atmosphere to decompose the binders, remove the solvents, andconvert the precursor to the desired ceramic matrix material. Due todecomposition of the binders, the result for each preform is a porousCMC fired body that may undergo densification, e.g., melt infiltration(MI), to fill the porosity and yield the respective CMC component.

Specific processing techniques and parameters for the above process willdepend on the particular composition of the materials. For example,silicon CMC components may be formed from fibrous material that isinfiltrated with molten silicon, e.g., through a process typicallyreferred to as the Silcomp process. Another technique of manufacturingCMC components is the method known as the slurry cast melt infiltration(MI) process. In one method of manufacturing using the slurry cast MImethod, CMCs are produced by initially providing plies of balancedtwo-dimensional (2D) woven cloth comprising silicon carbide(SiC)-containing fibers, having two weave directions at substantially90° angles to each other, with substantially the same number of fibersrunning in both directions of the weave. The term “siliconcarbide-containing fiber” refers to a fiber having a composition thatincludes silicon carbide, and preferably is substantially siliconcarbide. For instance, the fiber may have a silicon carbide coresurrounded with carbon, or in the reverse, the fiber may have a carboncore surrounded by or encapsulated with silicon carbide.

Other techniques for forming CMC components include polymer infiltrationand pyrolysis (PIP) and oxide/oxide processes. In PIP processes, siliconcarbide fiber preforms are infiltrated with a preceramic polymer, suchas polysilazane and then heat treated to form a SiC matrix. Inoxide/oxide processing, aluminum or alumino-silicate fibers may bepre-impregnated and then laminated into a preselected geometry.Components may also be fabricated from a carbon fiber reinforced siliconcarbide matrix (C/SiC) CMC. The C/SiC processing includes a carbonfibrous preform laid up on a tool in the preselected geometry. Asutilized in the slurry cast method for SiC/SiC, the tool is made up ofgraphite material. The fibrous preform is supported by a tool during achemical vapor infiltration process at about 1200° C., whereby the C/SiCCMC component is formed. In still other embodiments, 2D, 2.5D, and/or 3Dpreforms may be utilized in MI, CVI, PIP, or other processes. Forexample, cut layers of 2D woven fabrics may be stacked in alternatingweave directions as described above, or filaments may be wound orbraided and combined with 3D weaving, stitching, or needling to form2.5D or 3D preforms having multiaxial fiber architectures. Other ways offorming 2.5D or 3D preforms, e.g., using other weaving or braidingmethods or utilizing 2D fabrics, may be used as well.

Optionally, after processing, the CMC component (i.e., CMC airfoil 102,CMC inner band 104, and CMC outer band 106) may be finish machined, ifand as needed, and coated with one or more coatings, such as anenvironmental barrier coating (EBC). Moreover, the method describedabove is provided by way of example only. As an example, other knownmethods or techniques for curing composite plies, as well as fordensifying a CMC component, may be utilized. Alternatively, anycombinations of these or other known processes may be used. Further,although in the exemplary embodiments described herein the airfoilassembly 100 as a turbine nozzle fairing assembly, it will beappreciated that the airfoil assembly 100 could be any nozzle fairingassembly. Additionally, although described herein with respect to CMCairfoils 102, CMC inner bands 104, and CMC outer bands 106, the presentsubject matter could be applied to an airfoil assembly 100 formed fromany suitable material, such as a polymer matrix composite (PMC) materialor other composite material. Moreover, the airfoil assembly 100 maycomprise any appropriate combination of materials, e.g., in someembodiments, at least one of the airfoil 102, inner band 104, or outerband 106 may be formed from a different material than the remainder ofthe components of the airfoil assembly.

Accordingly, as described herein, an airfoil assembly having airfoilsseparate from an inner band and an outer band may be constructed. Byforming each component of the airfoil assembly as a separate piece,complicated CMC ply layups may be avoided, which may reduce thecomplexity of the manufacturing process for the airfoil assembly whileincreasing part yield and maintaining suitable compaction of eachcomponent part of the airfoil assembly. Further, separating each airfoilof the assembly from the inner and outer bands eliminates stresses dueto the thermal fight between the airfoil and bands. By eliminating suchstresses, the airfoil assemblies described herein are more defecttolerant than known integral airfoil and band designs, which opens upthe non-destructive examination limits of the assemblies. Separation ofthe airfoil and bands also enables doublet, triplet, etc. airfoilassemblies, which eliminate split line leakage. Therefore, compared toknown airfoil assembly designs, the airfoil assemblies described hereinhave improved producibility and reduced stresses, which reduces defectsand increases acceptability of the airfoil assemblies. Other advantagesof the subject matter described herein also may be realized by those ofordinary skill in the art.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal language of the claims.

What is claimed is:
 1. An airfoil assembly for a gas turbine engine,comprising: an airfoil having a concave pressure side opposite a convexsuction side and an inner end spaced apart from an outer end along aradial direction, the pressure side and the suction side extending alongan axial direction from a leading edge to a trailing edge, the airfoildefining a cavity, the airfoil including a trailing edge portion locatedaft of the cavity and defining the trailing edge, a first slot extendinglongitudinally along the radial direction defined in the trailing edgeportion for receipt of a first pin; and an inner band defining an inneropening shaped complementary to the inner end of the airfoil, whereinthe inner end of the airfoil is received with the inner opening, andwherein each of the inner band and airfoil are formed from a ceramicmatrix composite material.
 2. The airfoil assembly of claim 1, wherein asecond slot extending along the radial direction is defined in theairfoil for receipt of a second pin.
 3. The airfoil assembly of claim 2,wherein the first slot is defined in the inner end of the airfoil andthe second slot is defined in the outer end of the airfoil.
 4. Theairfoil assembly of claim 1, wherein the inner band includes a flowpathsurface, a non-flowpath surface opposite the flowpath surface, and afirst inner flange and a second inner flange extending along the radialdirection from the non-flowpath surface.
 5. The airfoil assembly ofclaim 4, wherein the inner band is secured to an inner support structureby a first inner fastener extending through the first inner flange and asecond inner fastener extending through the second inner flange.
 6. Theairfoil assembly of claim 5, wherein the first inner flange and thesecond inner flange extend from the first non-flowpath surface of theinner band such that the first inner flange and the second inner flangestraddle the inner support structure.
 7. The airfoil assembly of claim1, further comprising: a strut extending along the radial directionthrough a cavity defined by the airfoil; and a first pad defined at afirst radial location within the cavity and a second pad defined at asecond radial location within the cavity, the first radial locationdifferent from the second radial location.
 8. An airfoil assembly for agas turbine engine, comprising: an airfoil having a concave pressureside opposite a convex suction side and an inner end spaced apart froman outer end along a radial direction, the pressure side and the suctionside extending along an axial direction from a leading edge to atrailing edge, the airfoil defining a cavity, the airfoil including atrailing edge portion located aft of the cavity and defining thetrailing edge, a first slot extending longitudinally along the radialdirection defined in the trailing edge portion for receipt of a firstpin; and an outer band defining an outer opening shaped complementary tothe outer end of the airfoil, wherein the outer end of the airfoil isreceived with the outer opening, and wherein each of the outer band andairfoil are formed from a ceramic matrix composite material.
 9. Theairfoil assembly of claim 8, wherein a second slot extending along theradial direction is defined in the airfoil for receipt of a second pin.10. The airfoil assembly of claim 9, wherein the first slot is definedin the outer end of the airfoil and the second slot is defined in theinner end of the airfoil.
 11. The airfoil assembly of claim 8, whereinthe outer band includes a flowpath surface, a non-flowpath surfaceopposite the flowpath surface, and a first outer flange and a secondouter flange extending along the radial direction from the non-flowpathsurface.
 12. The airfoil assembly of claim 11, wherein the outer band issecured to an outer support structure by a first outer fastenerextending through the first outer flange and a second outer fastenerextending through the second outer flange.
 13. The airfoil assembly ofclaim 12, wherein the first outer flange and the second outer flangeextend from the second non-flowpath surface of the outer band such thatthe first outer flange and the second outer flange straddle the outersupport structure.
 14. The airfoil assembly of claim 8, furthercomprising: an inner band defining an inner opening shaped complementaryto the inner end of the airfoil, wherein the inner end of the airfoil isreceived with the inner opening, and wherein the inner band is formedfrom a ceramic matrix composite material.
 15. The airfoil assembly ofclaim 8, further comprising: a strut extending along the radialdirection through a cavity defined by the airfoil; and a first paddefined at a first radial location within the cavity and a second paddefined at a second radial location within the cavity, the first radiallocation different from the second radial location.
 16. An airfoilassembly for a gas turbine engine, comprising: an airfoil having aconcave pressure side opposite a convex suction side and an inner endspaced apart from an outer end along a radial direction, the pressureside and the suction side extending along an axial direction from aleading edge to a trailing edge, the airfoil defining a cavity, theairfoil including a trailing edge portion located aft of the cavity anddefining the trailing edge; an inner band defining an inner openingshaped complementary to the inner end of the airfoil; and an outer banddefining an outer opening shaped complementary to the outer end of theairfoil, wherein the inner end of the airfoil is received with the inneropening and the outer end of the airfoil is received within the outeropening, wherein the airfoil is constrained along the axial direction bya pair of pins, each pin of the pair of pins having a diameter and anaxial length, the axial length of each pin of the pair of pins extendingalong the radial direction into the airfoil, and wherein each of theinner band, outer band, and airfoil are formed from a ceramic matrixcomposite material.
 17. The airfoil assembly of claim 16, wherein a slotextending longitudinally along the radial direction is defined in thetrailing edge portion for receipt of one pin of the pair of pins. 18.The airfoil assembly of claim 16, wherein the inner band includes afirst flowpath surface, a first non-flowpath surface opposite the firstflowpath surface, and a first inner flange and a second inner flangeextending along the radial direction from the first non-flowpath surfaceadjacent a forward end of the inner band, wherein the outer bandincludes a second flowpath surface, a second non-flowpath surfaceopposite the second flowpath surface, and a first outer flange and asecond outer flange extending along the radial direction from the secondnon-flowpath surface, wherein the first outer flange extends from thesecond non-flowpath surface of the outer band adjacent an aft end of theouter band, and wherein the second outer flange extends from the secondnon-flowpath surface of the outer band adjacent a forward end of theouter band.
 19. The airfoil assembly of claim 11, wherein the inner bandincludes a first flowpath surface, a first non-flowpath surface oppositethe first flowpath surface, and a first inner flange and a second innerflange extending along the radial direction from the first non-flowpathsurface such that the first inner flange and the second inner flangestraddle the inner support structure, and wherein the outer bandincludes a second flowpath surface, a second non-flowpath surfaceopposite the second flowpath surface, and a first outer flange and asecond outer flange extending along the radial direction from the secondnon-flowpath surface such that the first outer flange and the secondouter flange straddle the outer support structure.
 20. The airfoilassembly of claim 16, wherein a first slot extending longitudinallyalong the radial direction is defined in the trailing edge portion forreceipt of a first pin of the pair of pins, wherein a second slotextending longitudinally along the radial direction is defined in theairfoil for receipt of a second pin of the pair of pins, wherein thefirst slot is defined in the outer end of the airfoil and the secondslot is defined in the inner end of the airfoil, wherein a firstaperture is defined in the outer support structure such that the firstpin passes through the first aperture into the first slot, and wherein asecond aperture is defined in the inner support structure such that thesecond pin passes through the second aperture into the second slot.